US9085372B2 - Aircraft comprising at least one engine having contra-rotating rotors - Google Patents
Aircraft comprising at least one engine having contra-rotating rotors Download PDFInfo
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- US9085372B2 US9085372B2 US13/055,305 US200913055305A US9085372B2 US 9085372 B2 US9085372 B2 US 9085372B2 US 200913055305 A US200913055305 A US 200913055305A US 9085372 B2 US9085372 B2 US 9085372B2
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Classifications
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D31/00—Power plant control systems; Arrangement of power plant control systems in aircraft
- B64D31/02—Initiating means
- B64D31/06—Initiating means actuated automatically
- B64D31/12—Initiating means actuated automatically for equalising or synchronising power plants
Definitions
- the invention relates to aircraft comprising at least one engine having contra-rotating rotors.
- It may be an engine with contra-rotating propellers such as a propfan type turboprop, or an engine of a rotary-wing aircraft such as a helicopter.
- FIG. 1 we will first describe the unbalance problem for a single rotating disc.
- This figure shows a disc forming a propeller 2 comprising blades 4 , in this case eight.
- the propeller can rotate freely around an axis 6 corresponding to its main geometric axis of symmetry.
- the propeller has a balancing fault such that the center of gravity of the propeller is not on axis 6 but is shifted radially from it.
- This center of gravity 8 is for example located on one of the blades 4 , as shown, rather exaggerated, on FIG. 1 .
- the propeller is rotated around its axis 6 in the direction shown by the arrow 10 .
- the center of gravity 8 therefore generates an unbalance force 12 exerted on the propeller on the axis 6 in the plane of the disc along a radial direction towards the outside and passing through point 8 .
- This force rotates in direction 10 .
- It is an inertial unbalance. Consequently, for any rotating disc whose center of inertia does not coincide with the center of rotation, an inertial unbalance produces a radial force in the plane of the disc as shown on FIG. 1 .
- the moving disc comprises bearing surfaces such as the faces of propeller blades.
- a setting fault or a shape fault on the bearing surfaces may therefore generate an aerodynamic unbalance.
- the aerodynamic unbalance force is exerted at a point 14 located away from axis 6 .
- the unbalance force is composed firstly of a traction force increment referenced 16 on FIG. 2 and located outside the plane of the propeller disc, and a drag force increment 18 located in the plane of the propeller disc.
- the unbalance of a rotating disc is therefore represented by the measured acceleration R1 in terms of amplitude (gain) and phase ( ⁇ ) in the axis of the fixed supporting structure at the machine speed of rotation ⁇ o as shown on FIG. 3 .
- This figure shows on a first curve 20 the graph of gain (in m/s ⁇ 2 ) against speed of rotation ⁇ (in rad/s), and on the second curve 22 the graph of phase ⁇ (in radians) against this speed.
- the following measurement method called the vector influence coefficient method, can be used.
- the initial acceleration R1 which represents the result of the action of the unbalance required
- unbalance masses of known weight are added to the rotating system to measure their effect on the measured acceleration. For example, an unbalance of unit mass is added to the disc at phase angle 0° and a new acceleration R2 (gain and phase) at speed ⁇ o is measured.
- inertial and aerodynamic unbalances may have to be measured separately.
- the above-mentioned technique can be used to do this, providing in addition that modifications of the speed of rotation and independent modifications of the torque request can be made, in order to distinguish between the source of unbalance due to inertia and the source of unbalance due to the aerodynamic characteristics of the rotor.
- unbalance diagnostic software programs supply balancing solution vectors which include one solution vector for the first rotor and one solution vector for the second rotor.
- Each solution vector includes a modulus and a phase angle. This operation will be carried out to characterize the inertial unbalance and then to characterize the aerodynamic unbalance.
- FIG. 5 shows these forces PROP1 24 and PROP2 26 which are exerted at the axis of rotation 6 common to the two discs.
- the two discs rotate in different directions, indicated respectively 28 and 30 on FIG. 5 .
- the unbalance forces 24 and 26 also rotate in opposite directions, respectively 28 and 30 .
- phase ⁇ is equal to 0 when the two radial forces PROP1 (or R disc1 ) and PROP2 (or R disc2 ) are in phase
- the moment located outside the plane is expressed at the center of the disc 1 for example.
- the lever arm is the axial distance between the planes of discs 1 and 2 .
- the direction of the major axis of the ellipse depends on the relative phase between the unbalance forces PROP1 and PROP2. For example, if the positions of the two discs are such that the two forces are in phase in the vertical axis, the maximum excitation in the plane of the discs will be directed vertically. In contrast, if the positions of the two discs are such that the two forces are in phase opposition (180° shift) in the vertical direction, the maximum excitation in the plane of the discs will be directed horizontally.
- One objective of the invention is to reduce the vibrations generated in the supporting structure by engines of this type.
- the invention therefore provides for an aircraft comprising at least one engine having contra-rotating rotors, the engine or at least one of the engines having unbalances associated with at least one ellipse, the aircraft comprising means capable of controlling the engine or at least one of the engines such that, at a given engine speed, the major axis of the ellipse or at least one of the ellipses extends in a direction for which vibrations generated by the engine or engines have a minimum intensity in at least one predetermined site, particularly in a predetermined area, of the aircraft.
- the vibrations produced by the engine or each engine are therefore mainly oriented in a direction in which they will be felt as little as possible at the site, for example from the fuselage. Consequently, we do not try to reduce them at source.
- the vibrations felt are therefore reduced simply and at low energy cost.
- the invention takes into account the characteristics of the engines, especially through the choice of direction of the major axis of the ellipse. It does not involve adding extra weights (e.g. tuned vibration absorbers) like the known filtering or dissipation techniques, nor introduction into the engine of specific external energy to implement an active function.
- the vibrations can be minimized at a single site, several sites at a time (using several respective sensors, for example) or in an entire zone corresponding for example to a complete component of the aircraft such as the fuselage.
- control means are able to control the engine or at least one of the engines such that the major axis of the ellipse or at least one of the ellipses extends in a direction of lower vibration transmissivity in a structure supporting the engine.
- the vibrations are channeled in this direction. They can therefore be dampened very efficiently.
- the overall vibrations generated by at least two of the engines or the two engines are therefore processed together to minimize their intensity.
- the site is specific to the engine and the control means are able to control the other engine such that the other engine is associated with at least one other predetermined site at which the vibrations generated by this other engine itself have minimum intensity.
- control means are able to control the engine or at least one of the engines such that the unbalance forces of the rotors are exerted parallel to an inertia plane of a structure supporting the engine, for example a main inertia plane or the inertia plane which has the largest inertia modulus.
- control means are able to control the engines such that unbalance forces and/or unbalance moments of the respective engines exerted on the aircraft compensate each other at least partially.
- the invention also provides for a method for controlling an aircraft comprising at least one engine having contra-rotating rotors, the engine or at least one of the engines having unbalances associated with at least one ellipse, wherein the engine or at least one of the engines is controlled such that, at a given engine speed, the major axis of the ellipse or at least one of the ellipses extends in a direction for which vibrations generated by the engine or engines have a minimum intensity in at least one predetermined site, particularly in a predetermined area, of the aircraft.
- the invention also provides for a computer program which includes instructions that can control execution of a method according to the invention when it is executed on a computer, and a data storage medium which includes such program.
- FIGS. 1 and 2 are front and perspective views respectively of a rotating disc of the prior art
- FIG. 3 shows a graph of the gain and phase of an unbalance force as a function of the speed of the rotating disc of FIG. 1 ;
- FIGS. 4 to 6 are diagrams showing, in vector form, unbalance forces in the disc of FIG. 1 ;
- FIGS. 7 and 8 are two diagrammatic elevation views of an aircraft showing two embodiments of the invention respectively.
- FIGS. 9 , 10 and 11 are diagrams showing different embodiments of the architecture for controlling the engines in the aircraft of FIGS. 7 and 8 .
- the aircraft 50 comprises a main structure 52 including in particular a fuselage 54 .
- the aircraft comprises engines 56 , in this case two, arranged symmetrically with respect to a median vertical plane of the fuselage 54 .
- each engine 56 is a propfan type turboprop comprising two contra-rotating propellers respectively referenced 58 a and 58 b , shown on FIG. 10 .
- the two propellers extend coaxially with reference to a common axis 60 , the front propeller 58 a extending in front of the rear propeller 58 b .
- the two propellers each have blades 61 . They can rotate in respectively opposite directions.
- Each engine 56 is connected to the fuselage 54 via a respective mast (or pylon) 62 .
- Direction Y is parallel to the average fiber of the mast 62 and goes through it.
- Direction X is horizontal and parallel to the longitudinal direction of the fuselage 54 and direction Z is perpendicular to directions X and Y.
- the longitudinal axes of the fuselage and of the engines are parallel. These axes may be different for an industrial application. In this case, either reference can be chosen.
- the inertial unbalance vibrations appearing in the engine 56 are not all transmitted to the mast in the same way. Depending on the orientation of the unbalance forces along axes X, Y and Z, the vibrations will be more or less well transmitted by the mast to the fuselage 54 .
- the synchrophasing of the two rotors of the engine 56 is controlled such that the major axis 68 of the ellipse of unbalances of this engine extends in direction Z.
- the vibrations generated by the engine 56 have minimum intensity since they have been poorly transmitted by the mast.
- the same transmissivity direction analysis is carried out on the other engine 56 and the latter controlled in a similar way such that the major axis of the corresponding ellipse is oriented in the direction in which the vibrations are the least well transmitted from this other engine to the fuselage 54 .
- each engine is processed independently of the other in order to minimize the vibrations that each one transmits to the fuselage.
- a sensor such as an accelerometer 165 placed at a predetermined site of the aircraft are taken into account.
- this sensor is located at the end of the mast 62 adjacent to the fuselage 54 . This sensor is used to determine the direction of the major axis 68 for which the vibrations measured by the sensor are lowest.
- This determination can be carried out during preliminary tests on the engine when the aircraft is stationary. In a variant, it can be carried out when the aircraft is in flight, for example during commercial service. In another variant, at least one first determination is carried out during preliminary tests on the engine when the aircraft is stationary, then at least one other determination is carried out when the aircraft is in flight, for example during commercial service, for example to refresh the determination of the best orientation.
- two angles 01 and 02 are identified, indicating the inclination of the major axis of the ellipse of the respective engines with respect to a fixed direction of the supporting structure, for example the vertical direction. These two angles could quite easily have different values.
- the major axis associated with the engine shown on the left of the figure is more inclined with respect to the vertical than that of the engine shown on the right.
- a sensor 267 is positioned at a predetermined location such as a critical area of the fuselage, for example at the lower part of the cabin designed to take passenger seats.
- a sensor 267 is positioned at a predetermined location such as a critical area of the fuselage, for example at the lower part of the cabin designed to take passenger seats.
- the values of 01 and 02 may be different from each other, including in absolute value. The best combination of angles 01 and 02 is therefore selected.
- the major axes of the ellipses are oriented in these preferred inclinations.
- the angles 01 and 02 can be determined during preliminary tests on the engine when the aircraft is stationary. In a variant, it can be determined when the aircraft is in flight, for example during commercial service. In another variant, at least one first determination is carried out during preliminary tests on the engine when the aircraft is stationary, then at least one other determination is carried out when the aircraft is in flight, for example during commercial service, for example to refresh the determination.
- a combination of angles 01 and is chosen for which those of the unbalance forces generating impacts, i.e. the oscillating forces, are oriented parallel to one of the main inertia planes of the supporting structure, for example in this plane.
- the inertia planes will be for example planes of symmetry of the supporting structure. The latter will be the entire aircraft or the structure supporting the engine locally, such as the mast. The impact of the vibrations appearing in the two engines will therefore be significantly reduced.
- the structure is in fact relatively insensitive to the vibrations exerted in such a plane or parallel to this plane.
- Several combinations of the angles 01 and 02 may be suitable.
- the plane of lowest inertia of the mast (which could for example include the direction of low transmissivity) could also be chosen. In this case, the vibrations will be minimized in the rest of the aircraft.
- the direction of lowest transmissivity could in fact be determined using the determination of this plane of lowest inertia of the mast.
- the inertia plane with the largest inertia modulus could be chosen. This will be the case for example in aircraft whose propulsion system is supported by a radial mast.
- the major axes 68 of the ellipses of the engines 156 are oriented such that the impact forces and the moments of the two engines act along the same line or along the same lines but in opposite directions on the two engines so that they compensate each other as seen from the supporting structure, in particular from the fuselage 54 .
- This embodiment will be adapted when the aircraft comprises several propulsion systems installed symmetrically on the aircraft structure.
- FIG. 9 shows the two contra-rotating rotors 58 a and 58 b of one of the engines 56 and 156 and the principle of an architecture for controlling the synchrophasing for each of the above-mentioned embodiments.
- the front rotor 58 a is the master rotor. It is controlled such that its speed and power have the values required by the engine operation.
- the rear rotor 58 b is in this case considered to be the slave rotor. Its speed of rotation is continuously adjusted to reduce the difference between its speed of rotation and that of the master rotor.
- the phase difference between the two rotors i.e. the orientation of the major axis 68 of the ellipse, is selected as indicated above in order to reduce the vibrations.
- the invention may also be implemented when the master rotor is the rear rotor and the front rotor is slave.
- the respective speeds of rotation of the two rotors are therefore measured with tachometers 71 attached to the two rotors.
- the information recorded by the tachometers is transmitted to a speed control module 73 which calculates the difference in speeds of rotation by calculating the difference between the inputs supplied by the two tachometers.
- the module 73 may include, as in this case, a suitable dynamic compensation algorithm in order to generate a corrective signal which is transmitted to a signal power amplifier 75 .
- the signal leaving the latter is transmitted to a unit 77 controlling the pitch of the rear rotor 58 b , which controls this rotor.
- the control may consist of a flow of fuel or a pitch control system, or both.
- the speed of the rear rotor is therefore adjusted so that it is as close as possible to that of the front rotor.
- the difference in phase angle between the two rotors is then adjusted to set the orientation of the major axis of the ellipse with respect to the fixed coordinate system, for example the coordinate system related to the supporting structure.
- a sensor 83 sensitive to the vibrations is therefore placed near the engine and the vibration intensity values it records are transmitted to a sampling and time-delay module 79 which transmits them to module 73 so that they can also be taken into account.
- data from the tachometer 31 associated with the rear slave rotor 58 b is transmitted to a pulse generator 81 also connected to the module 79 .
- the speeds of rotation of the rotors and their phase difference are therefore measured and adjusted.
- the angle controlling the optimum direction of the major axis of the ellipse must now be determined. This angle is calculated by the module 73 . Calculation is carried out using the principles described above. Synchrophasing is then implemented, taking this value into account.
- the moment located outside the plane and generated by the radial forces and the lever arm between the two discs can be used to compensate the moment generated outside the plane by the aerodynamic unbalance (due to the forces parallel to the axis exerted on the two discs at a point offset with respect to their center of rotation).
- FIG. 10 shows an embodiment of the control architecture used in the case of the invention to control a single engine, whether the aircraft comprises this engine alone or whether it comprises two engines controlled individually.
- Each engine is associated with mast accelerometers 74 , distributed on the mast 62 .
- mast accelerometers 74 there are three accelerometers, two near a front edge of the mast and one near the rear edge.
- the aircraft comprises control means 76 formed in this case by an EEC (Electronic Engine Controller) type unit.
- the unit 76 is connected by lines 80 to traditional components of the propellers by which the unit receives information concerning the position in degrees and speed in revolutions per minute of the propeller.
- the unit 76 also receives via respective lines 82 , 84 , 86 and 88 data concerning the air, the air sampling request, the thrust request and the aircraft at cruising speed.
- the aircraft comprises a synchrophasing unit 90 forming in this case the module 73 shown on FIG. 9 . It is connected by lines 92 to the three respective accelerometers 74 to receive acceleration data from them indicative of the vibrations running through the mast 62 .
- the unit 90 is also connected by a line 94 to the control unit 76 to receive position and speed data concerning each propeller 58 a and 58 b .
- a line 96 allows the synchrophasing unit to transmit speed setpoints for the front and rear propellers to the control means.
- the unit 76 can control a flow in a fuel circuit to adjust the phase of each propeller.
- the unit 76 therefore takes into account the data received to control the fuel flow in the engine and the pitch of the front and rear propellers, via the line 77 connecting it to the engine.
- the units 76 and 90 each form microprocessor electronic control means comprising at least one program stored on a data storage medium such as a hard disk or flash memory. At least one of the programs includes code instructions capable of controlling implementation of all or part of the method according to the invention when it is executed on these means forming a computer.
- Lines 82 , 84 , 86 , 88 , 94 and 96 are part of the aircraft onboard data communication network. It is for example an AFDX (Avionics Full Duplex Switch) type network.
- FIG. 11 shows an embodiment of the control architecture of two engines of an aircraft according to the invention when they are controlled together. Since the components are the same as those shown on FIG. 10 , they will not be described again.
- the synchrophasing unit 90 is common to the two engines 156 , whereas each engine has its own control unit 76 which receives the data already described and which has not all been shown.
- the two lines 94 of the respective engines lead to the unit 90 , while the two lines 96 of the respective engines leave it to lead to the respective units 76 .
- the unit 90 is responsible for synchrophasing the two engines according to the invention and transmits the necessary commands to the control units 76 .
- the invention confines the vibrations in the area of the source producing them.
- the invention takes advantage of the natural intrinsic behavior of the aircraft structure to dampen the vibrations generated by the engines.
- the invention aims to optimize the direction of the vibrations and to orient the vibrations in the preferred direction to minimize the effect, especially within the fuselage.
- the result is that, overall, the main supporting structure absorbs less energy than in the prior art. Dampening is effective and implements a small number of sensors.
- the direction of lowest transmissivity could be different from direction Z, to which it corresponds on FIG. 7 . It will depend on the aircraft architecture.
- the invention is applicable to engines comprising more than two rotating parts independent from each other.
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Abstract
Description
-
- the original unbalance b1 causes acceleration R1,
- the set (b1+b2) forming the sum of the original unbalance and of the unit unbalance causes an acceleration R2,
- by deduction, the unit unbalance b2 therefore generates the acceleration R2-R1. Concerning this subject, we refer to
FIG. 4 which shows in an orthonormal coordinate system the vectors R1, R2 and R2−R1 which have respectively phases φR1, φR2 and φ(R2−R1).
R(ωt)=R disc1(ωt)+R disc2(ωt)
R(ωt+π/2)=R disc1(ωt+π/2)−R disc2(ωt+π/2)
R(ωt+π)=−[R disc1(ωt)+R disc2(ωt)]
R(ωt+3 π/2)=R disc2(ωt+3 π/2)−R disc1(ωt+3 π/2)
M(ωt)=0
M(ωt+π/2)=[R disc1(ωt)+R disc2(ωt)]*leverarm
M(ωt+π)=0
M(ωt+3 π/2)=−[R disc1(ωt)+R disc2(ωt)]*leverarm
Claims (10)
Applications Claiming Priority (3)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
FR0855062A FR2934246B1 (en) | 2008-07-24 | 2008-07-24 | AIRCRAFT COMPRISING AT LEAST ONE ENGINE WITH CONTRAROTATIVE ROTORS |
FR0855062 | 2008-07-24 | ||
PCT/FR2009/051463 WO2010010292A2 (en) | 2008-07-24 | 2009-07-21 | Aircraft including at least one engine having counter-rotating rotors |
Publications (2)
Publication Number | Publication Date |
---|---|
US20110198440A1 US20110198440A1 (en) | 2011-08-18 |
US9085372B2 true US9085372B2 (en) | 2015-07-21 |
Family
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Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
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US13/055,305 Expired - Fee Related US9085372B2 (en) | 2008-07-24 | 2009-07-21 | Aircraft comprising at least one engine having contra-rotating rotors |
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US (1) | US9085372B2 (en) |
EP (1) | EP2303694B1 (en) |
JP (1) | JP2011528645A (en) |
CN (1) | CN102159463B (en) |
BR (1) | BRPI0916290A2 (en) |
CA (1) | CA2731960C (en) |
FR (1) | FR2934246B1 (en) |
RU (1) | RU2011106754A (en) |
WO (1) | WO2010010292A2 (en) |
Cited By (4)
Publication number | Priority date | Publication date | Assignee | Title |
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US20140180657A1 (en) * | 2011-08-17 | 2014-06-26 | Snecma | Method of determining the performance of at least one propeller of a turbomachine in an air stream under gyration |
US20170233066A1 (en) * | 2014-08-28 | 2017-08-17 | Sikorsky Aircraft Corporation | Pitch control system |
US10556699B2 (en) * | 2016-04-28 | 2020-02-11 | Airbus Operations Sas | Aircraft engine assembly comprising a pylon leading edge incorporated with an annular row of unfaired after-guide vanes |
US10676184B2 (en) | 2014-08-28 | 2020-06-09 | Sikorsky Aircraft Corporation | Pitch control system for an aircraft |
Families Citing this family (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
FR2934245B1 (en) * | 2008-07-24 | 2010-10-01 | Airbus France | AIRCRAFT COMPRISING A SYNCHROPHASED CONTROL ENGINE |
FR2934246B1 (en) * | 2008-07-24 | 2010-09-10 | Airbus France | AIRCRAFT COMPRISING AT LEAST ONE ENGINE WITH CONTRAROTATIVE ROTORS |
US10013900B2 (en) * | 2014-09-23 | 2018-07-03 | Amazon Technologies, Inc. | Vehicle noise control and communication |
US10395446B2 (en) * | 2016-05-04 | 2019-08-27 | Tecat Performance Systems, Llc | Integrated wireless data system for avionics performance indication |
US10745110B2 (en) * | 2018-06-29 | 2020-08-18 | Pratt & Whitney Canada Corp. | Propeller blade synchrophasing using phonic wheel |
US11203420B2 (en) * | 2019-05-03 | 2021-12-21 | Pratt & Whitney Canada Corp. | System and method for controlling engine speed in multi-engine aircraft |
US11613371B1 (en) | 2021-11-19 | 2023-03-28 | John Daniel Romo | Electric vacuum jet engine |
EP4266139A1 (en) * | 2022-04-19 | 2023-10-25 | Bühler AG | Monitoring machines |
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2008
- 2008-07-24 FR FR0855062A patent/FR2934246B1/en not_active Expired - Fee Related
-
2009
- 2009-07-21 WO PCT/FR2009/051463 patent/WO2010010292A2/en active Application Filing
- 2009-07-21 US US13/055,305 patent/US9085372B2/en not_active Expired - Fee Related
- 2009-07-21 JP JP2011519218A patent/JP2011528645A/en not_active Withdrawn
- 2009-07-21 CN CN200980136394.0A patent/CN102159463B/en not_active Expired - Fee Related
- 2009-07-21 BR BRPI0916290A patent/BRPI0916290A2/en not_active IP Right Cessation
- 2009-07-21 EP EP09740351A patent/EP2303694B1/en not_active Not-in-force
- 2009-07-21 CA CA2731960A patent/CA2731960C/en not_active Expired - Fee Related
- 2009-07-21 RU RU2011106754/11A patent/RU2011106754A/en unknown
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US10556699B2 (en) * | 2016-04-28 | 2020-02-11 | Airbus Operations Sas | Aircraft engine assembly comprising a pylon leading edge incorporated with an annular row of unfaired after-guide vanes |
Also Published As
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CN102159463B (en) | 2013-08-07 |
WO2010010292A2 (en) | 2010-01-28 |
FR2934246B1 (en) | 2010-09-10 |
BRPI0916290A2 (en) | 2016-07-19 |
JP2011528645A (en) | 2011-11-24 |
CA2731960A1 (en) | 2010-01-28 |
CA2731960C (en) | 2016-08-09 |
CN102159463A (en) | 2011-08-17 |
RU2011106754A (en) | 2012-08-27 |
FR2934246A1 (en) | 2010-01-29 |
EP2303694B1 (en) | 2013-01-30 |
US20110198440A1 (en) | 2011-08-18 |
EP2303694A2 (en) | 2011-04-06 |
WO2010010292A3 (en) | 2010-03-18 |
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